1. Field of the Invention
The invention relates to an engine turbine, and in particular to an assembly for controlling the flows of cooling air through the bottoms of the blade root sockets in the discs of a turbine rotor.
2. Summary of the Prior Art
The turbines of an aircraft engine generally comprise several stages of blades some of which have the blade root sockets arranged to be cooled by a flow of air taken from the compressor, and in order to reduce the amount of air taken from the compressor the cooling is effected in cascade fashion. Thus, only the first stage is cooled by the air coming directly from the compressor, the following stages being cooled by the air having already circulated in the preceding stage or stages.
The flow of air taken is designed to effect cooling under the severest operating conditions, such as during take-off, when there is experienced the highest engine loading and the greatest thermal gradients.
During the other phases of flight: i.e. cruising, idling, and approach, some stages can do without cooling. Accordingly, attempts have been made to achieve automatic control of the cooling air flows by making use of the dimensional variations of certain elements of the turbine when subjected to the variations in temperature of the gas flow passing through the blades.
For example, U.S. Pat. No. 3,975,901 discloses a device for controlling the clearance between a sealing ring and the ends of the rotor blades by cooling the ring with a gas flow of which the temperature and the rate of flow may be varied. This gas flow is obtained by mixing cold air taken from the compressor with hot gases taken from the exhaust duct from the combustion chamber. the cold air and the hot gases flow into two separate chambers each of which is provided in a radial wall with a circular row of openings. A disc co-operates with the radial wall and has two rows of openings of which at least one co-operates with one of the rows of openings in the radial wall. The disc also has, on its inside edge, a shoulder which is disposed adjacent the flow of the combustion gases passing through the blades. Temperature variations of the combustion gases produce dimensional variations of the ring which bring the rows of openings in the disc into complete or partial registry with the rows of openings of the chambers, thereby allowing a greater or lesser flow of cold air and hot gases into a mixing chamber which supplies the ring cooling circuit.
GB Pat. No. 1 484 288 is also concerned with a device for controlling the clearances of a sealing ring, the device utilising the intake of air from the high pressure compressor into two concentric chambers formed in the sealing ring. The ring is held by a resilient support inside a concentric member which is of relatively substantial mass and which forms, with the outer wall of the ring, a chamber which is also supplied by air from the high pressure compressor, this air being supplied through an annular passage which is closable in response to differential expansion of the member and of the ring.
The present invention is based on the same principle of controlling the intake of cooling air by varying closure of air passages by the expansion of an element subjected to the temperature variations of the gas flow passing through the blades of the turbine.